1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine rotor blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with multiple rows or stages or stator vanes that guide a high temperature gas flow through adjacent rotors of rotor blades to produce mechanical power and drive a bypass fan, in the case of an aero engine, or an electric generator, in the case of an IGT. In both cases, the turbine is also used to drive the compressor.
In the turbine section of the gas turbine engine, stages or rotor blades and stator vanes are used to guide the hot gas flow through and react with the rotor blades to drive the engine, to improve engine efficiency, the upstream stages of these airfoils (vanes and blades) are cooled with cooling air to produce convection cooling, impingement cooling, and even film cooling of the outer airfoil surfaces in order to allow for exposure to higher gas flow temperatures. The higher the turbine inlet temperature of the turbine, the higher will be the turbine efficiency and thus the engine efficiency. However, the highest temperature allowed is dependent upon the material properties of these airfoils, especially for the first stage airfoils, and the amount of cooling provided.
Higher levels of cooling can be used for these airfoils. However, since the pressurized cooling air is from the compressor, the more cooling air used from the compressor the more compressed air and work performed by the compressor that is not turned into useful work by the engine, the engine efficiency also decreases due to the extra work performed on compressing the cooling air which is then discharged into the hot gas flow so that not work is performed.
Especially for an industrial gas turbine engine, erosion or corrosion damage to a stator vane or a rotor blade in the turbine section can cause significant decrease in the engine performance or even an airfoil damaged so much that the engine must be prematurely shut down and the damaged airfoil replaced. An industrial gas turbine engine of the kind used in electric power production is intended to operate without stopping for a period of 40,000 hours or more. If an airfoil is damaged enough, the performance of the engine can be decreased such that the operating cost will be much higher. Thus, turbine airfoils are designed to minimize or eliminate the occurrence of hot spots that can result in these types of damage.
FIG. 1 shows a first stage turbine rotor blade of the prior art as disclosed in U.S. Pat. No. 5,947,687 issued to Mori et al. on Sep. 7, 1999 and entitled GAS TURBINE MOVING BLADE. FIG. 2 shows a cross section view of the cooling circuit for the prior art rotor blade of FIG. 1. This particular rotor blade includes two separate serpentine flow cooling circuit to provide cooling for the blade, one serpentine circuit in the forward section and another in the rear or aft section. The forward serpentine circuit is a 3-pass forward flowing serpentine with a first leg 11 for supplying pressurized cooling air, a second leg 12 and a third leg 13 located adjacent to a leading edge impingement channel 15 that includes a showerhead arrangement of film cooling holes 16 and even gill holes 17 on the pressure wall side. A row of film cooling holes 18 on the suction side wall is connected to the third leg 13.
The FIG. 2 blade includes a second serpentine circuit with a first leg 21 adjacent to the first leg 11 of the forward section serpentine circuit, a second leg 22 and a third leg 23 located adjacent to a trailing edge region of the blade airfoil. The first leg 21 is the supply channel for the aft serpentine flow circuit. The trailing edge section includes an arrangement of impingement holes or pedestals 25 that cool the trailing edge region, and then a row of exit holes or slots 26 to discharge the cooling air from the rear section serpentine flow circuit. the last leg 23 is connected through a row of impingement holes 28 to the trailing edge cooling circuit, and a row of film cooling holes 27 on the pressure side wall is connected to the last leg 23.
FIG. 3 shows a cross sectional side view of the blade cooling circuit for FIGS. 1 and 2 of the Mori et al invention. As seen in FIG. 3, the last leg 13 of the forward section serpentine circuit is located adjacent to the leading edge impingement channel 15, while the last leg 23 of the rear section serpentine flow circuit is located adjacent to the trailing edge cooling circuit with the pedestals or impingement holes 25. the cooling air flew in the rear section serpentine circuit flows from the root and into the first leg 21 toward the blade tip, then turns and flows into the second leg 22 toward the root section, and then turns again at the root and platform sections to flow up and into the third or last leg 23 along the trailing edge region. The cooling air from the trailing edge cooling circuit is discharged through the trailing edge exit slots 26 arranged along the trailing edge as seen in FIG. 4. A trailing edge fillet 31 forms a smooth transition between the airfoil surface 32 and the platform on the blade 33. In this prior art blade, a large amount of metal mass 34 is formed between the trailing edge fillet and the platform that is under-cooled in this arrangement.
For the blade trailing edge root section in the FIG. 4 prior art blade, due to the hot gas migration from the blade upper span down to the trailing edge and platform region, the blade aft fillet region experiences a hotter gas flow temperature. In addition, at the blade trailing edge fillet location 31, a higher heat transfer coefficient or heat load is formed on the fillet location due to the trailing edge wake effect. On top of a higher heat load onto the airfoil root section fillet 31, due to a stress concentration issue, the cooling slot for the airfoil trailing edge root section cannot be located low enough into the blade root section region to provide proper convection cooling. Cooling for this part of the airfoil trailing edge fillet region becomes especially difficult. A high thermally induced stress is predicted at the junction of the blade trailing edge and the platform location in the prior art blade design. Also, due to the different effectiveness levels of cooling used for the blade and the platform, and because of the mass metal distribution between the blade airfoil and the platform, the thermally induced strain during transient cycle (stopping and starting the engine, or going from steady state operation to less than steady state and back) becomes much more severe.